Aircraft with active aerosurfaces

ABSTRACT

An aircraft operable to transition between a forward flight mode and a vertical takeoff and landing flight mode. The aircraft includes an airframe having first and second wings. A plurality of propulsion assemblies is attached to the airframe with each of the propulsion assemblies including a nacelle and a tail assembly having at least one active aerosurface. A flight control system is operable to independently control each of the propulsion assemblies. For each of the propulsion assemblies, the tail assembly is rotatable relative to the nacelle such that the active aerosurface has a first orientation generally parallel to the wings and a second orientation generally perpendicular to the wings.

TECHNICAL FIELD OF THE DISCLOSURE

The present disclosure relates, in general, to aircraft operable totransition between a forward flight mode and a vertical takeoff andlanding flight mode and, in particular, to VTOL aircraft with activeaerosurfaces that are operable to selectively provide pitch control andyaw control to the aircraft.

BACKGROUND

Fixed-wing aircraft, such as airplanes, are capable of flight usingwings that generate lift responsive to the forward airspeed of theaircraft, which is generated by thrust from one or more jet engines orpropellers. The wings generally have an airfoil cross section thatdeflects air downward as the aircraft moves forward, generating the liftforce to support the airplane in flight. Fixed-wing aircraft, however,typically require a runway that is hundreds or thousands of feet longfor takeoff and landing.

Unlike fixed-wing aircraft, vertical takeoff and landing (VTOL) aircraftdo not require runways. Instead, VTOL aircraft are capable of takingoff, hover and landing vertically. One example of VTOL aircraft is ahelicopter which is a rotorcraft having one or more rotors that providelift and thrust to the aircraft. The rotors not only enable hover andvertical takeoff and landing, but also enable, forward, backward andlateral flight. These attributes make helicopters highly versatile foruse in congested, isolated or remote areas where fixed-wing aircraft maybe unable to takeoff and land. Helicopters, however, typically lack theforward airspeed of fixed-wing aircraft.

A tiltrotor aircraft is another example of a VTOL aircraft. Tiltrotoraircraft generate lift and propulsion using proprotors that aretypically coupled to nacelles mounted near the ends of a fixed wing. Thenacelles rotate relative to the fixed wing such that the proprotors havea generally horizontal plane of rotation for vertical takeoff, hover andlanding and a generally vertical plane of rotation for forward flight,wherein the fixed wing provides lift and the proprotors provide forwardthrust. In this manner, tiltrotor aircraft combine the vertical liftcapability of a helicopter with the speed and range of fixed-wingaircraft. Tiltrotor aircraft, however, typically suffer from downwashinefficiencies during vertical takeoff and landing due to interferencecaused by the fixed wing.

A further example of a VTOL aircraft is a tiltwing aircraft thatfeatures a rotatable wing that is generally horizontal for forwardflight and rotates to a generally vertical orientation for verticaltakeoff and landing. Propellers are coupled to the rotating wing toprovide the required vertical thrust for takeoff and landing and therequired forward thrust to generate lift from the wing during forwardflight. The tiltwing design enables the slipstream from the propellersto strike the wing on its smallest dimension, thus improving verticalthrust efficiency as compared to tiltrotor aircraft. Tiltwing aircraft,however, are more difficult to control during hover as the verticallytilted wing provides a large surface area for crosswinds typicallyrequiring tiltwing aircraft to have either cyclic rotor control or anadditional thrust station to generate a moment.

SUMMARY

In a first aspect, the present disclosure is directed to an aircraftoperable to transition between a forward flight mode and a verticaltakeoff and landing flight mode. The aircraft includes an airframehaving first and second wings. A plurality of propulsion assemblies isattached to the airframe with each of the propulsion assembliesincluding a nacelle and a tail assembly having at least one activeaerosurface. A flight control system is operable to independentlycontrol each of the propulsion assemblies. For each of the propulsionassemblies, the tail assembly is rotatable relative to the nacelle suchthat the active aerosurface has a first orientation generally parallelto the wings and a second orientation generally perpendicular to thewings.

In some embodiments, each of the propulsion assemblies may include anactuator that is operable to rotate the tail assembly relative to thenacelle. In certain embodiments, each of the propulsion assemblies mayinclude an actuator that is operable to translate the tail assemblyrelative to the nacelle between a retracted configuration and anextended configuration. In some embodiments, in the first orientation,each of the active aerosurfaces may be a horizontal stabilizer and/or anelevator to provide pitch control to the aircraft. In certainembodiments, in the second orientation, each of the active aerosurfacesmay be a vertical stabilizer and/or a rudder to provide yaw control tothe aircraft.

In some embodiments, during vertical takeoff and landing flightmaneuvers, each of the active aerosurfaces may be in the firstorientation. In certain embodiments, during hover flight maneuvers, eachof the active aerosurfaces may be in the first orientation. In someembodiments, during transitions from vertical takeoff and landing flightmaneuvers to forward flight maneuvers, each of the active aerosurfacesmay be in the first orientation. In certain embodiments, duringtransitions from forward flight maneuvers to vertical takeoff andlanding flight maneuvers, each of the active aerosurfaces may be in thefirst orientation. In some embodiments, during forward flight maneuvers,each of the active aerosurfaces may be in the second orientation.

In a second aspect, the present disclosure is directed to an aircraftoperable to transition between a forward flight mode and a verticaltakeoff and landing flight mode. The aircraft includes an airframehaving first and second wings. A plurality of propulsion assemblies isattached to the airframe with each of the propulsion assembliesincluding a nacelle, a tail assembly having at least one activeaerosurface, a first actuator and a second actuator. A flight controlsystem is operable to independently control each of the propulsionassemblies. The flight control system includes an active aerosurfacecontrol system operable to control operations of each of the first andsecond actuators. For each of the propulsion assemblies, the firstactuator is operable to rotate the tail assembly relative to the nacellesuch that the active aerosurface has a first orientation generallyparallel to the wings and a second orientation generally perpendicularto the wings. Also, for each of the propulsion assemblies, the secondactuator is operable to translate the tail assembly relative to thenacelle between a retracted configuration and an extended configuration.

In some embodiments, in the first orientation, each of the activeaerosurfaces may be a horizontal stabilizer and/or an elevator toprovide pitch control to the aircraft. In certain embodiments, in thesecond orientation, each of the active aerosurfaces may be a verticalstabilizer and/or a rudder to provide yaw control to the aircraft. Insome embodiments, during vertical takeoff and landing flight maneuvers,during hover flight maneuvers and/or during transitions between forwardflight maneuvers and vertical takeoff and landing flight maneuvers, eachof the active aerosurfaces may be in the first orientation, while duringforward flight maneuvers, each of the active aerosurfaces may be in thesecond orientation.

BRIEF DESCRIPTION OF THE DRAWINGS

For a more complete understanding of the features and advantages of thepresent disclosure, reference is now made to the detailed descriptionalong with the accompanying figures in which corresponding numerals inthe different figures refer to corresponding parts and in which:

FIGS. 1A-1F are schematic illustrations of an aircraft in accordancewith embodiments of the present disclosure;

FIGS. 2A-2B are block diagrams of propulsion assemblies for an aircraftin accordance with embodiments of the present disclosure;

FIGS. 3A-3I are schematic illustrations of an aircraft in a sequentialflight operating scenario in accordance with embodiments of the presentdisclosure;

FIGS. 4A-4D are schematic illustrations of a propulsion assembly for anaircraft in accordance with embodiments of the present disclosure;

FIGS. 5A-5H are schematic illustrations of an aircraft in various groundmaneuver configurations in accordance with embodiments of the presentdisclosure;

FIGS. 6A-6D are schematic illustrations of a propulsion assembly for anaircraft depicting the tilting degree of freedom and thrust vectorgeneration of a rotor assembly in accordance with embodiments of thepresent disclosure;

FIGS. 7A-7D are various views depicting the connections betweenpropulsion assemblies and an airframe of an aircraft in accordance withembodiments of the present disclosure;

FIG. 8 is a block diagram of an aircraft control system in accordancewith embodiments of the present disclosure;

FIGS. 9A-9B are isometric and exploded views of a rotor assemblyoperable to generate a variable thrust output and a variable thrustvector at a constant rotational speed for an aircraft in accordance withembodiments of the present disclosure; and

FIGS. 10A-10B are isometric and exploded views of a rotor assemblyoperable to generate a variable thrust output and a variable thrustvector at a constant rotational speed for an aircraft in accordance withembodiments of the present disclosure.

DETAILED DESCRIPTION

While the making and using of various embodiments of the presentdisclosure are discussed in detail below, it should be appreciated thatthe present disclosure provides many applicable inventive concepts,which can be embodied in a wide variety of specific contexts. Thespecific embodiments discussed herein are merely illustrative and do notdelimit the scope of the present disclosure. In the interest of clarity,not all features of an actual implementation may be described in thepresent disclosure. It will of course be appreciated that in thedevelopment of any such actual embodiment, numerousimplementation-specific decisions must be made to achieve thedeveloper's specific goals, such as compliance with system-related andbusiness-related constraints, which will vary from one implementation toanother. Moreover, it will be appreciated that such a development effortmight be complex and time-consuming but would be a routine undertakingfor those of ordinary skill in the art having the benefit of thisdisclosure.

In the specification, reference may be made to the spatial relationshipsbetween various components and to the spatial orientation of variousaspects of components as the devices are depicted in the attacheddrawings. However, as will be recognized by those skilled in the artafter a complete reading of the present disclosure, the devices,members, apparatuses, and the like described herein may be positioned inany desired orientation. Thus, the use of terms such as “above,”“below,” “upper,” “lower” or other like terms to describe a spatialrelationship between various components or to describe the spatialorientation of aspects of such components should be understood todescribe a relative relationship between the components or a spatialorientation of aspects of such components, respectively, as the devicedescribed herein may be oriented in any desired direction.

Referring to FIGS. 1A-1F in the drawings, various views of an aircraft10 having a versatile propulsion system are depicted. In the illustratedembodiment, aircraft 10 including an airframe 12 having wings 14, 16each have an airfoil cross-section that generates lift responsive to theforward airspeed of aircraft 10. Wings 14, 16 may be formed as singlemembers or may be formed from multiple wing sections. The outer skinsfor wings 14, 16 are preferably formed from high strength andlightweight materials such as fiberglass fabric, carbon fabric,fiberglass tape, carbon tape and combinations thereof that may be formedby curing together a plurality of material layers.

Extending generally perpendicularly between wings 14, 16 are trussstructures depicted as pylons 18, 20. Pylons 18, 20 are preferablyformed from high strength and lightweight materials such as fiberglassfabric, carbon fabric, fiberglass tape, carbon tape and combinationsthereof that may be formed by curing together a plurality of materiallayers. Preferably, wings 14, 16 and pylons 18, 20 are securablyattached together at the respective intersections by bolting, bondingand/or other suitable technique such that airframe 12 becomes a unitarymember. As illustrated, wings 14, 16 are polyhedral wings with wing 14having anhedral sections 14 a, 14 b and with wing 16 having dihedralsections 16 a, 16 b, as best seen in FIG. 1B. In this design, any liquidfuel stored in the anhedral sections 14 a, 14 b or dihedral sections 16a, 16 b of wings 14, 16 will gravity feed to sump locations. Wings 14,16 preferably include central passageways operable to contain energysources and communication lines. For example, as best seen in FIG. 1A,wing 14 includes energy sources 22 a, 22 b and communication lines 24 a,24 b. Energy sources 22 a, 22 b may be liquid fuel, batteries or othersuitable energy sources. In the case of liquid fuel energy sources 22 a,22 b, communication lines 24 a, 24 b form a fluid distribution networkwherein communication lines 24 a, 24 b access energy sources 22 a, 22 bproximate the outboard end of wing 14 due to the gravity feed enabled byanhedral sections 14 a, 14 b. A pumping system may be used to move theliquid fuel from the gravity sumps to the desired use location. In thecase of liquid fuel energy sources in wing 16, the communication linesin wing 16 would access the liquid fuel energy sources proximate theinboard end of wing 16 due to the gravity feed enabled by dihedralsections 16 a, 16 b.

In the illustrated embodiment, the versatile propulsion system includesa plurality of interchangeably propulsion assemblies 26 a, 26 b, 26 c,26 d that are independently attachable to and detachable from airframe12. As illustrated, propulsion assemblies 26 a, 26 b, 26 c, 26 d arepositioned on airframe 12 in a close coupled configuration. Asillustrated, the versatile propulsion system includes four independentlyoperating propulsion assemblies 26 a, 26 b, 26 c, 26 d. It should benoted, however, that a versatile propulsion system of the presentdisclosure could have any number of independent propulsion assembliesincluding six, eight, twelve, sixteen or other number of independentpropulsion assemblies. In the illustrated embodiment, propulsionassemblies 26 a, 26 b are securably attached to airframe 12 in a highwing configuration and propulsion assemblies 26 c, 26 d are securablyattached to airframe 12 in a low wing configuration by bolting or othersuitable technique, as best seen in FIG. 1B. Preferably, each propulsionassembly 26 a, 26 b, 26 c, 26 d includes a nacelle 28 a, 28 b, 28 c, 28d, as best seen in FIG. 1F, that houses a power source, an engine ormotor, a drive system, a rotor hub, actuators and an electronics nodeincluding, for example, controllers, sensors and communications elementsas well as other components suitable for use in the operation of apropulsion assembly. Each propulsion assembly 26 a, 26 b, 26 c, 26 dalso includes a tail assembly 46 a, 46 b, 46 c, 46 d having an activeaerosurface 48 a, 48 b, 48 c, 48 d, as best seen in FIGS. 1E and 1F. Inaddition, each propulsion assembly 26 a, 26 b, 26 c, 26 d has a rotorassembly including the rotor hub having a plurality of grips such asspindle grips and a proprotor 38 a, 38 b, 38 c, 38 d depicted as havingthree rotor blades each of which is coupled to one of the spindle gripsof the respective rotor hub such that the rotor blades are operable torotate with the spindle grips about respective pitch change axes, asdiscussed herein.

Aircraft 10 has a liquid fuel flight mode, wherein energy is provided toeach of the propulsion assemblies from liquid fuel. For example, in thisconfiguration, each of the propulsion assemblies may be represented bypropulsion assembly 26 a of FIGS. 1A and 2A. As illustrated, propulsionassembly 26 a includes a nacelle 28 a, one or more fuel tanks 30 a, aninternal combustion engine 32 a, a drive system 34 a, a rotor hub 36 a,a proprotor 38 a and an electronics node 40 a. In the liquid fuel flightmode, the fuel tanks of the propulsion assemblies may be connected tothe fluid distribution network of the airframe and serve as feeder tanksfor the IC engines. Alternatively, the liquid fuel system may be adistributed system wherein liquid fuel for each propulsion assembly isfully self-contained within the fuel tanks positioned within thenacelles, in which case, the wet wing system described above may not berequired. The IC engines may be powered by gasoline, jet fuel, diesel orother suitable liquid fuel. The IC engines may be rotary engines such asdual rotor or tri rotor engines or other high power-to-weight ratioengines. The drive systems may include multistage transmissions operablefor reduction drive such that optimum engine rotation speed and optimumproprotor rotation speed are enabled. The drive systems may utilizehigh-grade roller chains, spur and bevel gears, v-belts, high strengthsynchronous belts or the like. As one example, the drive system may be atwo-stage cogged belt reducing transmission including a 3 to 1 reductionin combination with a 2 to 1 reduction resulting in a 6 to 1 reductionbetween the engine and the rotor hub.

Aircraft 10 also has an electric flight mode, wherein energy is providedto each of the propulsion assemblies from an electric power source. Forexample, in this configuration, each of the propulsion assemblies may berepresented by propulsion assembly 26 b of FIGS. 1A and 2B. Asillustrated, propulsion assembly 26 b includes a nacelle 28 b, one ormore batteries 30 b, an electric motor 32 b, a drive system 34 b, arotor hub 36 b, a proprotor 38 b and an electronics node 40 b. In theelectric flight mode, the electric motors of each propulsion assemblyare preferably operated responsive to electrical energy from the batteryor batteries disposed with that nacelle, thereby forming a distributedelectrical system. Alternatively or additionally, electrical power maybe supplied to the electric motors and/or the batteries disposed withthe nacelles from the energy sources, such as energy sources 22 a, 22 b,carried by airframe 12 via the communication lines, such ascommunication lines 24 a, 24 b.

Aircraft 10 also has a mixed flight mode, wherein energy is provided tosome of the propulsion assemblies from an electric power source andenergy is provided to other of the propulsion assemblies from liquidfuel. The mixed flight mode of aircraft 10 is evident from theillustrated embodiment of FIG. 1A. As another alternative, some or allof the engines of the propulsion assembly 26 a, 26 b, 26 c, 26 d may behydraulic motors operated responsive to a distributed hydraulic fluidsystem wherein high pressure hydraulic sources or generators are housedwithin each nacelle. Alternatively or additionally, a common hydraulicfluid system integral to or carried by airframe 12 may be used.

To transition aircraft 10 among liquid fuel flight mode, electric flightmode and mixed flight mode, propulsion assembly 26 a, 26 b, 26 c, 26 dare preferably standardized and interchangeable units that are mostpreferably line replaceable units enabling easy installation and removalfrom airframe 12, as discussed herein. In addition, regardless of thecurrent flight mode configuration of aircraft 10, the use of linereplaceable units is beneficial in maintenance situations if a fault isdiscovered with one of the propulsion assemblies. In this case, thefaulty propulsion assembly can be decoupled from airframe 12 by simpleoperations such as unbolting structural members, disconnecting power,data and/or fluid couplings and other suitable procedures. Anotherpropulsion assembly can then be attached to airframe 12 by connectingpower, data and/or fluid couplings, bolting structural members togetherand other suitable procedures.

The rotor assemblies of each propulsion assembly 26 a, 26 b, 26 c, 26 dare preferably lightweight, rigid members that may include swashyokemechanisms operable for collective pitch control and thrust vectoring.Proprotors 38 a, 38 b, 38 c, 38 d each include a plurality of proprotorblades that are securably attached to spindle grips of the respectiverotor hub. The blades are preferably operable for collective pitchcontrol and may additional be operable for cyclic pitch control. As analternative, the pitch of the blades may be fixed, in which case, thrustis determined by changes in the rotational velocity of the proprotors.In the illustrated embodiment, the rotor hubs have a tilting degree offreedom to enable thrust vectoring.

To accommodate the tilting degree of freedom of the rotor hubs, wings14, 16 have a unique swept wing design, which is referred to herein asan M-wing design. For example, as best seen in FIG. 1A, wing 14 hasswept forward portions 14 c, 14 d and swept back portions 14 e, 14 f.Propulsion assembly 26 a is coupled to a wing stanchion positionedbetween swept forward portion 14 c and swept back portion 14 e.Likewise, propulsion assembly 26 b is coupled to a wing stanchionpositioned between swept forward portion 14 d and swept back portion 14f. Wing 16 has a similar M-wing design with propulsion assemblies 26 c,26 d similarly coupled to wing stanchions positioned between sweptforward and swept back portions. In this configuration, each rotor hubis operable to pivot about a mast axis, such as mast axis 42 a and mastaxis 42 b, to control the direction of the thrust vector while avoidingany interference between proprotor 38 a, 38 b, 38 c, 38 d and wings 14,16. In the illustrated embodiment, the maximum angle of the thrustvector may preferably be between about 10 degrees and about 30 degrees,may more preferably be between about 15 degrees and about 25 degrees andmay most preferably be about 20 degrees. In embodiments having a maximumthrust vector angle of 20 degrees, the thrust vector may be resolved toany position within a 20-degree cone swung about the mast centerlineaxis. Notably, using a 20-degree thrust vector yields a lateralcomponent of thrust that is about 34 percent of total thrust. Eventhough the propulsion assemblies of the present disclosure have beendescribed as having certain nacelles, power sources, engines, drivesystems, rotor hubs, proprotors and tail assemblies, it is to beunderstood by those having ordinary skill in the art that propulsionassemblies having other components or combinations of componentssuitable for use in a versatile propulsion system are also possible andare considered to be within the scope of the present disclosure.

As best seen in FIGS. 1A, 1C and 1E, aircraft 10 includes landing geardepicted as including wheels 44 a, 44 b, 44 c, 44 d. The landing gearmay be passively operated pneumatic landing struts or actively operatedtelescoping landing struts disposed within tail assemblies 46 a, 46 b,46 c, 46 d of propulsion assemblies 26 a, 26 b, 26 c, 26 d. As discussedherein, wheels 44 a, 44 b, 44 c, 44 d enable aircraft 10 to taxi andperform other ground maneuvers. The landing gear may provide a passivebrake system or may include active brakes such as an electromechanicalbraking system or a manual braking system to facilitate parking asrequired during ground operations and/or passenger ingress and egress.In the illustrated embodiment, each tail assembly 46 a, 46 b, 46 c, 46 dincludes an active aerosurface 48 a, 48 b, 48 c, 48 d that is controlledby an active aerosurface control module of a flight control system 40.During various flight operations, active aerosurfaces 48 a, 48 b, 48 c,48 d may operate as vertical stabilizers, horizontal stabilizers,rudders and/or elevators to selectively provide pitch control and yawcontrol to aircraft 10.

As best seen in FIGS. 1A and 2A-2B, each of the propulsion assemblies 26a, 26 b, 26 c, 26 d includes one or more actuators, such as actuators 52a disposed within nacelle 28 a and actuators 52 b disposed withinnacelle 28 b, that are operable to change the orientation, theconfiguration and/or the position of the active aerosurface 48 a, 48 b,48 c, 48 d. For each propulsion assembly and responsive to commands fromthe active aerosurface control module of flight control system 40, afirst actuator is operable to rotate the tail assembly relative to thenacelle (see FIG. 4A). This operation also rotates the activeaerosurface relative to the nacelle such that the active aerosurfaces 48a, 48 b, 48 c, 48 d may be generally parallel to wings 14, 16, as bestseen in FIG. 1E, the active aerosurfaces 48 a, 48 b, 48 c, 48 d may begenerally perpendicular to wings 14, 16, as best seen in FIG. 1F, or theactive aerosurfaces 48 a, 48 b, 48 c, 48 d may be located in anyorientation therebetween. Active aerosurfaces 48 a, 48 b, 48 c, 48 d maybe fixed, in which case, when active aerosurfaces 48 a, 48 b, 48 c, 48 dare generally perpendicular to wings 14, 16 they may be referred to asvertical stabilizers that provide yaw control to aircraft 10. Similarly,when active aerosurfaces 48 a, 48 b, 48 c, 48 d are generally parallelto wings 14, 16 they may be referred to as horizontal stabilizers thatprovide pitch control to aircraft 10.

For each propulsion assembly and responsive to commands from the activeaerosurface control system of flight control module 40, a secondactuator may be operable to tilt all or a portion of each activeaerosurface relative to the tail assembly (see FIGS. 4C-4D). In thiscase, when active aerosurfaces 48 a, 48 b, 48 c, 48 d are generallyperpendicular to wings 14, 16 and are tilted relative to tail assemblies46 a, 46 b, 46 c, 46 d, active aerosurfaces 48 a, 48 b, 48 c, 48 d maybe referred to as rudders that provide yaw control to aircraft 10.Likewise, when active aerosurfaces 48 a, 48 b, 48 c, 48 d are generallyparallel to wings 14, 16 and are tilted relative to tail assemblies 46a, 46 b, 46 c, 46 d, active aerosurfaces 48 a, 48 b, 48 c, 48 d may bereferred to as elevators that provide pitch control to aircraft 10.

For each propulsion assembly and responsive to commands from the activeaerosurface control module of flight control system 40, a third actuatoris operable to translate the tail assembly relative to the nacellebetween a retracted configuration and an extended configuration (seeFIG. 4B). In the extended configuration of tail assemblies 46 a, 46 b,46 c, 46 d relative to nacelles 28 a, 28 b, 28 c, 28 d, as best seen inFIG. 1F, active aerosurfaces 48 a, 48 b, 48 c, 48 d are able to generatea greater moment, which can be beneficial in stabilizing aircraft 10. Itshould be noted that during certain flight maneuvers, it may bebeneficial to have active aerosurfaces 48 a, 48 b, 48 c, 48 d operatingas vertical stabilizers and/or rudders such as during forward flightmaneuvers, as best seen in FIG. 1F. Likewise, during other flightmaneuvers, it may be beneficial to have active aerosurfaces 48 a, 48 b,48 c, 48 d operating as horizontal stabilizers and/or elevators such asduring hover flight maneuvers, vertical takeoff and land flightmaneuvers and transition between forward flight and VTOL flight, as bestseen in FIG. 1E. In addition, it may be beneficial to have some ofactive aerosurfaces 48 a, 48 b, 48 c, 48 d operating as horizontalstabilizers and/or elevators and other of active aerosurfaces 48 a, 48b, 48 c, 48 d operating as vertical stabilizers and/or rudders duringcertain flight maneuvers. Further, it may be beneficial to have one ormore of active aerosurfaces 48 a, 48 b, 48 c, 48 d operating in aposition between the horizontal stabilizer position and the verticalstabilizer position during certain flight maneuvers.

Flight control system 40 of aircraft 10, such as a digital flightcontrol system, may be located within a central passageway of wing 14,as best seen in FIG. 1A. In the illustrated embodiment, flight controlsystem 40 is a triply redundant flight control system including threeindependent flight control computers. Use of triply redundant flightcontrol system 40 having redundant components improves the overallsafety and reliability of aircraft 10 in the event of a failure inflight control system 40. Flight control system 40 preferably includesnon-transitory computer readable storage media including a set ofcomputer instructions executable by one or more processors forcontrolling the operation of the versatile propulsion system. Flightcontrol system 40 may be implemented on one or more general-purposecomputers, special purpose computers or other machines with memory andprocessing capability. For example, flight control system 40 may includeone or more memory storage modules including, but is not limited to,internal storage memory such as random access memory, non-volatilememory such as read only memory, removable memory such as magneticstorage memory, optical storage, solid-state storage memory or othersuitable memory storage entity. Flight control system 40 may be amicroprocessor-based system operable to execute program code in the formof machine-executable instructions. In addition, flight control system40 may be selectively connectable to other computer systems via aproprietary encrypted network, a public encrypted network, the Internetor other suitable communication network that may include both wired andwireless connections.

Flight control system 40 communicates via a communications network 54with the electronics nodes of each propulsion assembly 26 a, 26 b, 26 c,26 d, such as electronics node 40 a of propulsion assembly 26 a andelectronics node 40 b of propulsion assembly 26 b, as best seen in FIGS.1A and 2A-2B. Flight control system 40 receives sensor data from andsends flight command information to the electronics nodes of eachpropulsion assembly 26 a, 26 b, 26 c, 26 d such that each propulsionassembly 26 a, 26 b, 26 c, 26 d may be individually and independentlycontrolled and operated. In both manned and unmanned missions, flightcontrol system 40 may autonomously control some or all aspects of flightoperation for aircraft 10. Flight control system 40 is also operable tocommunicate with remote systems, such as a transportation servicesprovider system via a wireless communications protocol. The remotesystem may be operable to receive flight data from and provide commandsto flight control system 40 to enable remote flight control over some orall aspects of flight operation for aircraft 10, in both manned andunmanned missions.

Aircraft 10 includes a pod assembly, illustrated as passenger podassembly 50, that is selectively attachable to airframe 12 betweenpylons 18, 20. In the illustrated embodiment, pylons 18, 20 includereceiving assemblies for coupling with pod assembly 50. Preferably, theconnection between pylons 18, 20 and pod assembly 50 allows pod assembly50 to rotate and translate relative to airframe 12 during flightoperations. In addition, one or more communication channels may beestablished between pod assembly 50 and airframe 12 when pod assembly 50is attached therewith. For example, a quick disconnect harness may becoupled between pod assembly 50 and airframe 12 to allow a pilot withinpod assembly 50 to receive flight data from and provide commands toflight control system 40 to enable onboard pilot control over some orall aspects of flight operation for aircraft 10.

As best seen in FIG. 1E, aircraft 10 is in a vertical takeoff andlanding mode. As illustrated, wings 14, 16 are generally above podassembly 50 with wing 14 forward of and wing 16 aft of pod assembly 50and with wings 14, 16 disposed in generally the same horizontal plane.As noted, flight control system 40 independently controls and operateseach propulsion assembly 26 a, 26 b, 26 c, 26 d. In one example, flightcontrol system 40 is operable to independently control collective pitchand adjust the thrust vector of each propulsion assembly 26 a, 26 b, 26c, 26 d, which can be beneficial in stabilizing aircraft 10 duringvertical takeoff, vertical landing and hover. As best seen in FIG. 1F,aircraft 10 is in a forward flight mode. Wings 14, 16 are generallyforward of pod assembly 50 with wing 14 below and wing 16 above podassembly 50 and with wings 14, 16 disposed in generally the samevertical plane. In the illustrated embodiment, the proprotor blades ofpropulsion assemblies 26 a, 26 d rotate counterclockwise while theproprotor blades of propulsion assemblies 26 b, 26 c rotate clockwise tobalance the torque of aircraft 10.

Referring next to FIGS. 3A-3I in the drawings, a sequentialflight-operating scenario of aircraft 10 is depicted. In the illustratedembodiment, passenger pod assembly 50 is attached to airframe 12. It isnoted, however, that passenger pod assembly 50 may be selectivelydisconnected from airframe 12 such that a single airframe can beoperably coupled to and decoupled from numerous passenger pod assembliesfor numerous missions over time. As best seen in FIG. 3A, aircraft 10 ispositioned on the ground with the tail assemblies 46 positioned in apassive brake configuration, as discussed herein. When aircraft 10 isready for a mission, flight control system 40 commences operations toprovide flight control to aircraft 10 which may be autonomous flightcontrol, remote flight control, onboard pilot flight control or anycombination thereof. For example, it may be desirable to utilize onboardpilot flight control during certain maneuvers such as takeoff andlanding but rely on remote or autonomous flight control during hover,forward flight and/or transitions between forward flight and VTOLoperations. As best seen in FIG. 3B, aircraft 10 is in its verticaltakeoff and landing mode and has lifted pod assembly 50 into the air.Preferably, each tail assembly 46 is rotated about the mast axis 56 ofthe respective nacelle 28 (see FIG. 4A) such that active aerosurfaces 48are generally parallel with wings 14, 16 to aid in stabilization duringhover and to be properly positioned to provide pitch control during thetransition to forward flight. After vertical assent to the desiredelevation, aircraft 10 may begin the transition from vertical takeoff toforward flight.

As best seen in FIGS. 3B-3D, as aircraft 10 transitions from verticaltakeoff and landing mode to forward flight mode, airframe 12 rotatesabout pod assembly 50 such that pod assembly 50 is maintained in agenerally horizontal attitude for the safety and comfort of passengers,crew and/or cargo carried in pod assembly 50. This is enabled by apassive and/or active connection between airframe 12 and pod assembly50. For example, a gimbal assembly may be utilized to allow passiveorientation of pod assembly 50 relative to airframe 12. This may beachieved due to the shape and the center of gravity of pod assembly 50wherein aerodynamic forces and gravity tend to bias pod assembly 50toward the generally horizontal attitude. Alternatively or additionally,a gear assembly, a clutch assembly or other suitably controllablerotating assembly may be utilized that allows for pilot controlled,remote controlled or autonomously controlled rotation of pod assembly 50relative to airframe 12 as aircraft 10 transitions from vertical takeoffand landing mode to forward flight mode. Preferably, as best seen inFIG. 3D, each tail assembly 46 is rotated about the mast axis 56 of therespective nacelle 28 (see FIG. 4A) such that active aerosurfaces 48 aregenerally perpendicular with wings 14, 16 to aid in stabilization andprovide yaw control during forward flight.

As best seen in FIGS. 3D-3E, once aircraft 10 has completed thetransition to forward flight mode, it may be desirable to adjust thecenter of gravity of aircraft 10 to improve its stability andefficiency. In the illustrated embodiment, this can be achieved byshifting pod assembly 50 forward relative to airframe 12 using an activeconnection between airframe 12 and pod assembly 50. For example,rotation of a gear assembly of pod assembly 50 relative to a rackassembly of airframe 12 or other suitable translation system may be usedto shift pod assembly 50 forward relative to airframe 12 under pilotcontrol, remote control or autonomous control. Preferably, as best seenin FIG. 3E, each tail assembly 46 is extended relative to the respectivenacelle 28 (see FIG. 4B) such that active aerosurfaces 48 providegreater stabilization and yaw control during forward flight.

When aircraft 10 begins its approaches to the destination, pod assembly50 is preferably returned to the aft position relative to airframe 12and each tail assembly 46 is preferably retracted (see FIG. 4B) androtated (see FIG. 4A) relative to the respective nacelle 28 such thatactive aerosurfaces 48 are generally parallel with wings 14, 16, as bestseen in FIG. 3F. Aircraft 10 may now begin its transition from forwardflight mode to vertical takeoff and landing mode. As best seen in FIGS.3F-3H, during the transition from forward flight mode to verticaltakeoff and landing flight mode, airframe 12 rotates about pod assembly50 such that pod assembly 50 is maintained in the generally horizontalattitude for the safety and comfort of passengers, crew and/or cargocarried in pod assembly 50. Once aircraft 10 has completed thetransition to vertical takeoff and landing flight mode, as best seen inFIG. 3H, aircraft 10 may commence its vertical descent to a surface.After landing, aircraft 10 may engage in ground maneuvers, if desired.Upon completion of any ground maneuvers, each tail assembly 46 isrotated about the mast axis 56 of the respective nacelle 28 (see FIG.4A) to return to the passive brake configuration, as best seen in FIG.3I.

Referring next to FIGS. 5A-5H and 6A-6D, the omnidirectional groundmaneuver capabilities of aircraft 10 will now be described. Asillustrated, aircraft 10 includes an airframe 12 having wings 14, 16. Aplurality of propulsion assemblies, referred to individually andcollectively as propulsion assembly 26 or propulsion assemblies 26, areattached to airframe 12. Two of the propulsion assemblies 26 are coupledto wing 14 and two of the propulsion assemblies 26 are coupled to wing16. In the illustrated embodiment, propulsion assemblies 26 aregenerally symmetrically positioned or disposed about an aircraftrotational axis 58 to provide stability to aircraft 10. Each of thepropulsion assemblies 26 includes a nacelle 28 having a mast axis 56, arotor assembly 60 having a tilting degree of freedom relative to mastaxis 56 and a tail assembly 46 rotatable about mast axis 56. It shouldbe noted that each rotor assembly 60 includes a rotor hub 36 and aproprotor 38, as best see in FIG. 1A. Each tail assembly 46 forms alanding gear including at least one wheel 44 having a rotational axis62. As discussed herein, a flight control system 40 is operable toindependently control each of the propulsion assemblies 26 includingtilting each rotor assembly 60 and rotating each tail assembly 46. Foreach propulsion assembly 20, flight control system 40 is operable totilt rotor assembly 60 in any direction relative to mast axis 56, asbest seen in FIGS. 6A-6D, enabling a thrust vector 64 to be resolvedwithin a thrust vector cone relative to mast axis 56. In the illustratedembodiment, thrust vector 64 has a maximum angle relative to mast axis56 of about twenty degrees. When propulsion assemblies 26 are beingoperated including rotation of rotor assemblies 60 about mast axis 56and tilting of rotor assemblies 60 relative to mast axis 56, the thrustvectors 64 generated by rotor assemblies 60 have a vertical component 66and a horizontal component 68. Importantly, the horizontal component 68of thrust vectors 64 is operable to provide the energy source requiredto roll aircraft 10 during ground maneuvers.

The omnidirectional ground maneuver capabilities of aircraft 10 areachieved by controlling thrust vectors 66 of rotor assemblies 60relative to rotational axes 62 of wheels 44 of tail assemblies 46. Forexample, for each propulsion assembly 26, rotor assembly 60 and tailassembly 46 have complementary configurations in which the horizontalcomponent 68 of thrust vector 64 is generally perpendicular torotational axis 62 of wheel 44. In this case, the thrust is pushingaircraft 10 in a direction that efficiently turns wheel 44. In addition,for each propulsion assembly 26, rotor assembly 60 and tail assembly 46have non complementary configurations in which the horizontal component68 of thrust vector 64 is not perpendicular to rotational axis 62 ofwheel 44. In this case, the thrust may be pushing aircraft 10 in adirection that inefficiently turns wheel 44 or in a direction that doesnot turn wheel 44.

In FIG. 5A, rotor assemblies 60 and tail assemblies 46 havecomplementary configurations in which the horizontal components 68 ofthrust vectors 64 are generally perpendicular to the respectiverotational axes 62 of wheels 44. Rotational axes 62 of wheels 44 aregenerally parallel with wings 14, 16 which creates a fore/aft groundmaneuver configuration, as indicated by the directional arrowintersecting aircraft rotational axis 58. In the illustrated embodiment,aircraft 10 would move in the forward direction due to the aftwarddirection of the horizontal components 68 of thrust vectors 64. Asdiscussed herein, wings 14, 16 are polyhedral wings having anhedral anddihedral sections. Thus, the use of terms such as “generally parallel,”“generally perpendicular” and “generally congruent” take into accountsuch angular variations.

In FIG. 5B, rotor assemblies 60 and tail assemblies 46 havecomplementary configurations. Rotational axes 62 of wheels 44 aregenerally perpendicular with wings 14, 16 which creates a lateral groundmaneuver configuration, as indicated by the directional arrowintersecting aircraft rotational axis 58. In the illustrated embodiment,aircraft 10 would move to its left, when viewed from the front, due tothe horizontal components 68 of thrust vectors 64 pointing to the right.FIGS. 5C and 5D illustrate tail assemblies 46 in two of a plurality ofradial ground maneuver configurations, as indicated by the directionalarrows intersecting respective aircraft rotational axes 58. Rotorassemblies 60 and tail assemblies 46 have complementary configurationsand each of the rotational axes 62 of wheels 44 has been rotated to thesame relative orientation between generally parallel and generallyperpendicular with wings 14, 16. In this manner, aircraft 10 is operablefor omnidirectional ground maneuvers.

FIG. 5E illustrates tail assemblies 46 in a turning ground maneuverconfiguration, as indicated by the directional arrow intersectingaircraft rotational axis 58. Aircraft 10 is being urged to the left,when viewed from the front, due to the horizontal components 68 ofthrust vectors 64 pointing to the right. As with all the previousexamples, rotor assemblies 60 have a common tilting configuration inwhich each of the rotor assemblies 60 is tilting in the same direction.In this case, however, some of the rotor assemblies 60 and tailassemblies 46 have complementary configurations while others of therotor assemblies 60 and tail assemblies 46 have non complementaryconfigurations. More specifically, as aircraft 10 moves to the left, therear rotor assemblies 60 and tail assemblies 46 have complementaryconfigurations in which rotational axes 62 of wheels 44 are generallyperpendicular with wings 14, 16. The front rotor assemblies 60 and tailassemblies 46 have non complementary configurations in which rotationalaxes 62 of wheels 44 have been rotated to a configuration betweengenerally parallel and generally perpendicular with wings 14, 16, thusproviding front wheel steering for aircraft 10.

FIG. 5F illustrates tail assemblies 46 in an alternate turning groundmaneuver configuration, as indicated by the directional arrowintersecting aircraft rotational axis 58. Aircraft 10 is being urgedforward due to the horizontal components 68 of thrust vectors 64pointing aftward. Rotor assemblies 60 have a common tiltingconfiguration in which each of the rotor assemblies 60 is tilting in thesame direction. The front rotor assemblies 60 and tail assemblies 46have complementary configurations in which rotational axes 62 of wheels44 are generally parallel with wings 14, 16. The rear rotor assemblies60 and tail assemblies 46 have non complementary configurations in whichrotational axes 62 of wheels 44 have been rotated to a configurationbetween generally parallel and generally perpendicular with wings 14,16, thus providing rear wheel steering for aircraft 10. Alternatively oradditionally, tail assemblies 46 may have a four wheel steering groundmaneuver configuration.

FIG. 5G illustrates tail assemblies 46 in a rotation ground maneuverconfiguration, as indicated by the directional arrow around aircraftrotational axis 58. Rotor assemblies 60 and tail assemblies 46 havecomplementary configurations. Rotor assemblies 60, however, have a noncommon tilting configuration in which rotor assemblies 60 are titling ina plurality of directions and, in the illustrated embodiment, are eachtilting in a different direction. In addition, rotational axis 62 ofeach wheel 44 intersects rotational axis 58 of aircraft 10. In thisrotation ground maneuver configuration of rotor assemblies 60 and tailassemblies 46, aircraft 10 will rotate about rotational axis 58 in acounter clockwise direction, when viewed from above.

FIG. 5H illustrates tail assemblies 46 in a passive brake configuration,as indicated by the stop symbol intersecting aircraft rotational axis58. Rotational axes 62 of adjacent wheels 44 are generally perpendicularwith each other. In addition, rotational axis 62 of each wheel 44 doesnot intersect rotational axis 58 of aircraft 10. Rotor assemblies 60 andtail assemblies 46 are illustrated in non complementary configurations,however, regardless of the relative configuration of rotor assemblies 60and tail assemblies 46, the horizontal components 68 of any thrustvectors 64 will not cause aircraft 10 to roll in any direction and willnot cause aircraft 10 to rotate about rotational axis 58 as the passivebrake configuration of tail assemblies 46 prevents such movement.

Referring to FIGS. 7A-7D in the drawings, the connections betweenpropulsion assemblies 26 a, 26 b, 26 c, 26 d and airframe 12 will now bediscussed. As illustrated, propulsion assemblies 26 a, 26 b, 26 c, 26 dare elements of the versatile propulsion system wherein, propulsionassemblies 26 a, 26 b, 26 c, 26 d are interchangeably attachable toairframe 12 as line replaceable units. Airframe 12 includes wings 14, 16and pylons 18, 20. Pod assembly 50 is supported between pylons 18, 20and is preferably rotatable and translatable relative to airframe 12. Asillustrated, wings 14, 16 each have an M-wing design. Wing 14 has sweptforward portions 14 c, 14 d and swept back portions 14 e, 14 f. Sweptforward portion 14 c and swept back portion 14 e meet at leading apex 14g. Swept forward portion 14 d and swept back portion 14 f meet atleading apex 14 h. Wing 16 has swept forward portions 16 c, 16 d andswept back portions 16 e, 16 f. Swept forward portion 16 c and sweptback portion 16 e meet at leading apex 16 g. Swept forward portion 16 dand swept back portion 16 f meet at leading apex 16 h. Each of the sweptforward portions 14 c, 14 d, 16 c, 16 d has a swept angle, which isdepicted relative to swept forward portion 16 d as angle 70. Each of theswept back portions 14 e, 14 f, 16 e, 16 f has a swept angle, which isdepicted relative to swept back portion 16 f as angle 72. Preferably,each of swept angles 70 are generally congruent with one another, eachof swept angles 72 are generally congruent with one another and sweptangles 70 are generally congruent with swept angles 72 such thatairframe 12 is symmetric. In the illustrated embodiment, swept angles70, 72 may preferably be between about 10 degrees and about 30 degrees,may more preferably be between about 15 degrees and about 25 degrees andmay most preferably be about 20 degrees.

Airframe 12 includes four stanchions 74, only stanchions 74 a, 74 bpositioned on the front of wing 14 being visible in FIG. 7A. Two similarstanchions are positioned aft of wing 16. In the illustrated embodiment,stanchion 74 a is located proximate leading apex 14 g and stanchion 74 bis located proximate leading apex 14 h. Similarly, the two stanchions ofwing 16 are located proximate leading apex 16 g and proximate leadingapex 16 h, respectively. Each of the stanchions 74 includes a flange 76having a bolt pattern, only flanges 76 a, 76 b being visible in FIG. 7A.As best seen in FIG. 7C, stanchion 74 a includes an interface panel 78 adepicted with two power sockets 80 a, four data or communication sockets82 a and two fluid sockets 84 a. As illustrated, sockets 80 a, 82 a, 84a are substantially flush or integrated with panel 78 a. As should beapparent to those having ordinary skill in the art, each stanchion ofthe present disclosure will include a similar panel with similarsockets. In addition, even though a particular arrangement of socketshas been depicted and described, those having ordinary skill in the artshould understand that the stanchions of the present disclosure couldhave other numbers of sockets in other arrangements.

Propulsion assemblies 26 a, 26 b, 26 c, 26 d each includes a flange 86,only flanges 86 c, 86 d being visible in FIG. 7A. Each flange 86 has abolt pattern that matches the bolt pattern of flanges 76 such thatpropulsion assemblies 26 can be interchangeably bolted to any one of thestanchions 74 to create a mechanical connection therebetween. As bestseen in FIG. 7D, propulsion assembly 26 a includes an interface panel 88a depicted with two power cables 90 a, four data or communication cables92 a and two fluid cables 94 a. Power cables 90 a are operable to couplewith power sockets 80 a to established electrical connections betweenairframe 12 and propulsion assembly 26 a. For example, these connectionsenable electrical power from batteries 30 b of airframe 12 to beprovided to components within propulsion assembly 26 a such aselectronics node 40 a, an electric motor and/or other electricalcomponents.

Communication cables 92 a are operable to couple with communicationsockets 82 a to established data communication between airframe 12 andpropulsion assembly 26 a. For example, these connections enable flightcontrol system 40 to communicate with electronics node 40 a to providecommand and control information to propulsion assembly 26 a and receivesensor and feedback information from propulsion assembly 26 a. Fluidcables 94 a are operable to couple with fluid sockets 84 a toestablished fluid communication between airframe 12 and propulsionassembly 26 a. For example, these connections enable liquid fuel fromairframe 12 to be provided to a fuel tank and/or an internal combustionengine of propulsion assembly 26 a. Alternatively or additionally, theseconnections may enable hydraulic fluid from airframe 12 to provide powerto hydraulic components within propulsion assembly 26 a. As should beapparent to those having ordinary skill in the art, each propulsionassembly of the present disclosure will include a similar panel withsimilar cables. In addition, even though a particular arrangement ofcables has been depicted and described, those having ordinary skill inthe art should understand that the propulsion assemblies of the presentdisclosure could have other numbers of cables in other arrangements thatpreferably mate with corresponding sockets of the stanchions of thepresent disclosure.

As illustrated, stanchions 74 provide standoff between propulsionassemblies 26 and wings 14, 16. By providing standoff between propulsionassemblies 26 and wings 14, 16, the aerodynamics of aircraft 10 areimproved by effectively creating more wing surface to provide liftduring various flight maneuvers. The M-wing design of wings 14, 16 andthe attachment of propulsion assemblies 26 to stanchions 74 proximateforward apexes 14 g, 14 h, 16 g, 16 h enables to rotor assemblies 60 tobe mounted in close proximity to forward apexes 14 g, 14 h, 16 g, 16 hand be tilted to generate variable thrust vectors 64. In the illustratedembodiment, each rotor assembly 60 has a tilting degree of freedomrelative to the respective mast axis 56 enabling thrust vectors 64 to beresolved within a thrust vector cone about mast axis 56, as best seen inFIGS. 6A-6D. The maximum angle of the thrust vector may preferably bebetween about 10 degrees and about 30 degrees, may more preferably bebetween about 15 degrees and about 25 degrees and may most preferably beabout 20 degrees. Importantly, the maximum thrust vector angle ispreferably generally congruent with swept angles 70, 72 of wings 14, 16,thereby avoiding interference between rotor assemblies 60 and airframe12.

Referring additionally to FIG. 8 in the drawings, a block diagramdepicts an aircraft control system 100 operable for use with aircraft 10of the present disclosure. In the illustrated embodiment, system 100includes three primary computer based subsystems; namely, an airframesystem 102, a passenger pod assembly system 104 and a remote system 106.As discussed herein, the aircraft of the present disclosure may beoperated autonomously responsive to commands generated by flight controlsystem 108 that preferably includes a non-transitory computer readablestorage medium including a set of computer instructions executable by aprocessor. Flight control system 108 may be a triply redundant systemimplemented on one or more general-purpose computers, special purposecomputers or other machines with memory and processing capability. Forexample, flight control system 108 may include one or more memorystorage modules including, but is not limited to, internal storagememory such as random access memory, non-volatile memory such as readonly memory, removable memory such as magnetic storage memory, opticalstorage, solid-state storage memory or other suitable memory storageentity. Flight control system 108 may be a microprocessor-based systemoperable to execute program code in the form of machine-executableinstructions. In addition, flight control system 108 may be selectivelyconnectable to other computer systems via a proprietary encryptednetwork, a public encrypted network, the Internet or other suitablecommunication network that may include both wired and wirelessconnections.

In the illustrated embodiment, flight control system 108 includes acommand module 110 and a monitoring module 112. It is to be understoodby those skilled in the art that these and other modules executed byflight control system 108 may be implemented in a variety of formsincluding hardware, software, firmware, special purpose processors andcombinations thereof. Flight control system 108 receives input from avariety of sources including internal sources such as sensors 114,controllers 116 and propulsion assemblies 118-122, and external sourcessuch as passenger pod assembly system 104, remote system 106 as well asglobal positioning system satellites or other location positioningsystems and the like. For example, flight control system 108 may receivea flight plan including starting and ending locations for a mission frompassenger pod assembly system 104 and/or remote system 106. Thereafter,flight control system 108 is operable to autonomously control allaspects of flight of an aircraft of the present disclosure.

For example, during the various operating modes of aircraft 10 includingvertical takeoff and landing flight mode, hover flight mode, forwardflight mode and transitions therebetween, command module 110 providescommands to controllers 116. These commands enable independent operationof each propulsion assembly 118-122 including tilting the rotorassemblies, adjusting the pitch of the proprotor blades, rotating,tilting and/or extending the tail assemblies and the like. Flightcontrol system 108 receives feedback from controllers 116 and eachpropulsion assembly 118-122. This feedback is processes by monitoringmodule 112 that can supply correction data and other information tocommand module 110 and/or controllers 116. Sensors 114, such aspositioning sensors, attitude sensors, speed sensors, environmentalsensors, fuel sensors, temperature sensors, location sensors and thelike also provide information to flight control system 108 to furtherenhance autonomous control capabilities.

Some or all of the autonomous control capability of flight controlsystem 108 can be augmented or supplanted by remote flight control from,for example, remotes system 106 such as a transportation servicesprovider. Remote system 106 may include one or computing systems thatmay be implemented on general-purpose computers, special purposecomputers or other machines with memory and processing capability. Forexample, the computing systems may include one or more memory storagemodules including, but is not limited to, internal storage memory suchas random access memory, non-volatile memory such as read only memory,removable memory such as magnetic storage memory, optical storagememory, solid-state storage memory or other suitable memory storageentity. The computing systems may be microprocessor-based systemsoperable to execute program code in the form of machine-executableinstructions. In addition, the computing systems may be connected toother computer systems via a proprietary encrypted network, a publicencrypted network, the Internet or other suitable communication networkthat may include both wired and wireless connections. The communicationnetwork may be a local area network, a wide area network, the Internet,or any other type of network that couples a plurality of computers toenable various modes of communication via network messages using assuitable communication techniques, such as transmission controlprotocol/internet protocol, file transfer protocol, hypertext transferprotocol, internet protocol security protocol, point-to-point tunnelingprotocol, secure sockets layer protocol or other suitable protocol.Remote system 106 communicates with flight control system 108 via acommunication link 124 that may include both wired and wirelessconnections.

Remote system 106 preferably includes one or more flight data displaydevices 126 configured to display information relating to one or moreaircraft of the present disclosure. Display devices 126 may beconfigured in any suitable form, including, for example, liquid crystaldisplays, light emitting diode displays, cathode ray tube displays orany suitable type of display. Remote system 106 may also include audiooutput and input devices such as a microphone, speakers and/or an audioport allowing an operator to communicate with, for example, a pilot onboard a pod assembly. The display device 126 may also serve as a remoteinput device 128 if a touch screen display implementation is used,however, other remote input devices, such as a keyboard or joystick, mayalternatively be used to allow an operator to provide control commandsto an aircraft being operated responsive to remote control.

Some or all of the autonomous and/or remote flight control of anaircraft of the present disclosure can be augmented or supplanted byonboard pilot flight control from an attached passenger pod assemblyincluding system 104. Passenger pod assembly system 104 preferablyincludes a non-transitory computer readable storage medium including aset of computer instructions executable by a processor and may beimplemented by a general-purpose computer, a special purpose computer orother machine with memory and processing capability. Passenger podassembly system 104 may include one or more memory storage modulesincluding, but is not limited to, internal storage memory such as randomaccess memory, non-volatile memory such as read only memory, removablememory such as magnetic storage memory, optical storage memory,solid-state storage memory or other suitable memory storage entity.Passenger pod assembly system 104 may be a microprocessor-based systemoperable to execute program code in the form of machine-executableinstructions. In addition, passenger pod assembly system 104 may beconnectable to other computer systems via a proprietary encryptednetwork, a public encrypted network, the Internet or other suitablecommunication network that may include both wired and wirelessconnections. Passenger pod assembly system 104 communicates with flightcontrol system 108 via a communication channel 130 that preferablyincludes a wired connection.

Passenger pod assembly system 104 preferably includes a cockpit displaydevice 132 configured to display information to an onboard pilot.Cockpit display device 132 may be configured in any suitable form,including, for example, as one or more display screens such as liquidcrystal displays, light emitting diode displays and the like or anyother suitable display type including, for example, a display panel, adashboard display, an augmented reality display or the like. Passengerpod assembly system 104 may also include audio output and input devicessuch as a microphone, speakers and/or an audio port allowing an onboardpilot to communicate with, for example, an operator of a remote system.Cockpit display device 132 may also serve as a pilot input device 134 ifa touch screen display implementation is used, however, other userinterface devices may alternatively be used to allow an onboard pilot toprovide control commands to an aircraft being operated responsive toonboard pilot control including, for example, a control panel,mechanical control devices or other control devices. As should beapparent to those having ordinarily skill in the art, through the use ofsystem 100, an aircraft of the present disclosure can be operatedresponsive to a flight control protocol including autonomous flightcontrol, remote flight control or onboard pilot flight control andcombinations thereof.

Referring next to FIGS. 9A-9B of the drawings, therein is depicted arotor assembly for use on an aircraft 10 that is operable to generate avariable thrust output and a variable thrust vector at a constantrotational speed and that is generally designated 200. In theillustrated embodiment, rotor assembly 200 includes a mast 202 that ispreferably rotated at a constant speed responsive to torque androtational energy provided by the engine and drive system of therespective propulsion assembly. Mast 202 rotates about a mast axis 204.A ball joint 206 is positioned about mast 202 but does not rotate withmast 202. Instead, in the illustrated embodiment, ball joint 206includes a flange 208 that is coupled to the airframe of aircraft 10 bybolting or other suitable technique. Ball joint 206 preferably has anouter spherical surface 210 that is operable to receive an innerspherical surface 211 of a tilt control assembly 212 thereon such thattilt control assembly 212 has a tilting degree of freedom relative toball joint 206. Tilt control assembly 212 does not rotate with mast 202.Instead, a tilting plate 214 of tilt control assembly 212 is coupled tothe airframe with a scissor mechanism 216 that includes a hinge 218 anda ball joint 220 that is received in a ball socket (not visible) oftilting plate 214. Two control rods 222 are coupled to tilting plate 214such that control rods 222 are operable to push and pull tilting plate214, thus actuating the tilting degree of freedom of tilt controlassembly 212. Control rods 222 may be electrically, hydraulically and/ormechanically controlled responsive to flight control commands receivedfrom flight control system 108 via autonomous flight control, remoteflight control, onboard pilot flight control or combinations thereof.

A rotor hub 224 is rotatably coupled to tilt control assembly 212 by abearing assembly depicted as a ball bearing assembly 226, which providesfor low friction relative rotation between rotor hub 224 and tiltcontrol assembly 212. Rotor hub 224 is rotated by mast 202 via arotational joint 228, such as a universal joint or a constant velocityjoint, that is coupled to mast 202 by bolting or other suitabletechnique. Rotational joint 228 provides a torque path between mast 202and rotor hub 224. Rotational joint 228 has a splined connector 230 thatis received within a splined portion 231 of a drive arm assembly 232 ofrotor hub 224. The splined mating surfaces allow rotor hub 224 totranslate relative to rotational joint 228 and thus mast 202 duringrotary operations. Rotor hub 224 rotates in a rotational plane aboutmast axis 204. The rotational plane may be normal to mast axis 204 whentilt control assembly 212 is not tilted relative to ball joint 206. Inaddition, the rotational plane may have an angle relative to mast axis204 when tilt control assembly 212 is tilted relative to ball joint 206as rotor hub 224 tilts with tilt control assembly 212 responsive toactuation of control rods 222. Rotor hub 224 including a plurality ofspindle grips depicted as three spindle grips 234 a, 234 b, 234 c, inthe illustrated embodiment. Spindle grips 234 a, 234 b, 234 c extendgenerally radially outwardly from the body of rotor hub 224. As bestseen in the exploded section, spindle grip 234 b includes a spindleassembly 234 d having outboard and inboard bearings 234 e, 234 f ontowhich grip assembly 234 g is secured and operable to rotate thereabout.Spindle grips 234 a, 234 c preferably have similar construction andoperation such that spindle grips 234 a, 234 b, 234 c are operable torotate about respective pitch change axes 236 a, 236 b, 236 c. In theillustrated embodiment, each spindle grip 234 a, 234 b, 234 c includes arespective pinion gear 238 a, 238 b, 238 c. A rotor blade (see forexample FIGS. 6A-6D) is coupled to each of the spindle grips 234 a, 234b, 234 c at respective devises 240 a, 240 b, 240 c by bolting, pinningor other suitable technique. The pitch of the rotor blades is controlledresponsive to rotation of the spindle grips 234 a, 234 b, 234 c aboutthe respective pitch change axes 236 a, 236 b, 236 c.

Rotor assembly 200 includes a collective pitch control mechanism 242. Inthe illustrated embodiment, collective pitch control mechanism 242includes a spider assembly 244 having a plurality of arms extendingtherefrom depicted as three arms 246 a, 246 b, 246 c. Each arm 246 a,246 b, 246 c includes a rack gear 248 a, 248 b, 248 c that is operableto mate with a respective pinion gear 238 a, 238 b, 238 c such thattranslation of spider assembly 244 relative to rotor hub 224 responsiveto operation of one or more actuators 250 rotates each spindle grip 234a, 234 b, 234 c about the respective pitch change axis 236 a, 236 b, 236c to collectively adjust the pitch of the rotor blades, therebygenerating the variable thrust output at a constant rotational speed. Inaddition, actuation of tilt control assembly 212 changes the rotationalplane of rotor hub 224 relative to mast axis 204, thereby generating thevariable thrust vector. As the control rods 222 are operable to tilttilting plate 214 in any direction relative to ball joint 206, therotational plane of rotor hub 224 may be tilted in any directionrelative to mast axis 204 thus enabling resolution of the thrust vectorwithin a thrust vector cone relative to mast axis 204. In someembodiments, the thrust vector cone may have a maximum angle relative tomast axis 204 of between about ten degrees and about thirty degrees. Inother embodiments, the thrust vector cone may have a maximum anglerelative to mast axis 204 of between about fifteen degrees and abouttwenty-five degrees. In additional embodiments, the thrust vector conemay have a maximum angle relative to mast axis 204 of about twentydegrees. It is noted that when rotor assemblies 200 are used in theversatile propulsion system of aircraft 10, each rotor assembly 200 ispositioned proximate a leading apex in the M-wing design thus enablingthe disclosed resolution of the thrust vector within the thrust vectorcone without interference between the rotor blades and the airframe.

Referring next to FIGS. 10A-10B of the drawings, therein is depicted arotor assembly for use on an aircraft 10 that is operable to generate avariable thrust output and a variable thrust vector at a constantrotational speed and that is generally designated 300. In theillustrated embodiment, rotor assembly 300 includes a mast 302 that ispreferably rotated at a constant speed responsive to torque androtational energy provided by the engine and drive system of therespective propulsion assembly. Mast 302 rotates about a mast axis 304.A ball joint 306 is positioned about mast 302 but does not rotate withmast 302. Instead, in the illustrated embodiment, ball joint 306includes a flange 308 that is coupled to the airframe of aircraft 10 bybolting or other suitable technique. Ball joint 306 preferably has anouter spherical surface 310 that is operable to receive an innerspherical surface 311 of a tilt control assembly 312 thereon such thattilt control assembly 312 has a tilting degree of freedom relative toball joint 306. Tilt control assembly 312 does not rotate with mast 302.Instead, a tilting plate 314 of tilt control assembly 312 is coupled tothe airframe with a scissor mechanism 316 that includes a hinge 318 anda ball joint 320 that is received in a ball socket (not visible) oftilting plate 314. Two control rods 322 are coupled to tilting plate 314such that control rods 322 are operable to push and pull tilting plate314, thus actuating the tilting degree of freedom of tilt controlassembly 312. Control rods 322 may be electrically, hydraulically and/ormechanically controlled responsive to flight control commands receivedfrom flight control system 108 via autonomous flight control, remoteflight control, onboard pilot flight control or combinations thereof.

A rotor hub 324 is rotatably coupled to tilt control assembly 312 by abearing assembly depicted as a ball bearing assembly 326, which providesfor low friction relative rotation between rotor hub 324 and tiltcontrol assembly 312. Rotor hub 324 is rotated by mast 302 via arotational joint 328, such as a universal joint or a constant velocityjoint, that is coupled to mast 302 by bolting or other suitabletechnique. Rotational joint 328 provides a torque path between mast 302and rotor hub 324. Rotational joint 328 has a splined connector 330 thatis received within a splined portion 331 of a drive arm assembly 332 ofrotor hub 324. The splined mating surfaces allow rotor hub 324 totranslate relative to rotational joint 328 and thus mast 302 duringrotary operations. Rotor hub 324 rotates in a rotational plane aboutmast axis 304. The rotational plane may be normal to mast axis 304 whentilt control assembly 312 is not tilted relative to ball joint 306. Inaddition, the rotational plane may have an angle relative to mast axis304 when tilt control assembly 312 is tilted relative to ball joint 306as rotor hub 324 tilts with tilt control assembly 312 responsive toactuation of control rods 322. Rotor hub 324 including a plurality ofspindle grips depicted as three spindle grips 334 a, 334 b, 334 c, inthe illustrated embodiment. Spindle grips 334 a, 334 b, 334 c extendgenerally radially outwardly from the body of rotor hub 324. As bestseen in the exploded section, spindle grip 334 b includes a spindleassembly 334 d having outboard and inboard bearings 334 e, 334 f ontowhich grip assembly 334 g is secured and operable to rotate thereabout.Spindle grips 334 a, 334 c preferably have similar construction andoperation such that spindle grips 334 a, 334 b, 334 c are operable torotate about respective pitch change axes 336 a, 336 b, 336 c. Asillustrated, each spindle grip 334 a, 334 b, 334 c includes a respectivepinion gear 338 a, 338 b, 338 c. A rotor blade (see for example FIGS.6A-6D) is coupled to each of the spindle grips 334 a, 334 b, 334 c atrespective devises 340 a, 340 b, 340 c by bolting, pinning or othersuitable technique. The pitch of the rotor blades is controlledresponsive to rotation of the spindle grips 334 a, 334 b, 334 c aboutthe respective pitch change axes 336 a, 336 b, 336 c.

Rotor assembly 300 includes a collective pitch control mechanism 342. Inthe illustrated embodiment, collective pitch control mechanism 342includes a ring assembly 344 having an input gear 346 and an output gear348 that is operable to mate with each of the pinion gears 338 a, 338 b,338 c such that rotation of ring assembly 344 relative to rotor hub 324responsive to operation of one or more actuators 350 rotates eachspindle grip 334 a, 334 b, 334 c about the respective pitch change axis336 a, 336 b, 336 c to collectively adjust the pitch of the rotorblades, thereby generating the variable thrust output at a constantrotational speed. In addition, actuation of tilt control assembly 312changes the rotational plane of rotor hub 324 relative to mast axis 304,thereby generating the variable thrust vector. As the control rods 322are operable to tilt tilting plate 314 in any direction relative to balljoint 306, the rotational plane of rotor hub 324 may be tilted in anydirection relative to mast axis 304 thus enabling resolution of thethrust vector within a thrust vector cone relative to mast axis 304. Insome embodiments, the thrust vector cone may have a maximum anglerelative to mast axis 304 of between about ten degrees and about thirtydegrees. In other embodiments, the thrust vector cone may have a maximumangle relative to mast axis 304 of between about fifteen degrees andabout twenty-five degrees. In additional embodiments, the thrust vectorcone may have a maximum angle relative to mast axis 304 of about twentydegrees. It is noted that when rotor assemblies 300 are used in theversatile propulsion system of aircraft 10, each of rotor assembly 300is positioned proximate a leading apex in the M-wing design thusenabling the thrust vector to be resolved within the thrust vector conewithout interference between the rotor blades and the airframe.

The foregoing description of embodiments of the disclosure has beenpresented for purposes of illustration and description. It is notintended to be exhaustive or to limit the disclosure to the precise formdisclosed, and modifications and variations are possible in light of theabove teachings or may be acquired from practice of the disclosure. Theembodiments were chosen and described in order to explain the principalsof the disclosure and its practical application to enable one skilled inthe art to utilize the disclosure in various embodiments and withvarious modifications as are suited to the particular use contemplated.Other substitutions, modifications, changes and omissions may be made inthe design, operating conditions and arrangement of the embodimentswithout departing from the scope of the present disclosure. Suchmodifications and combinations of the illustrative embodiments as wellas other embodiments will be apparent to persons skilled in the art uponreference to the description. It is, therefore, intended that theappended claims encompass any such modifications or embodiments.

What is claimed is:
 1. An aircraft operable to transition between aforward flight mode and a vertical takeoff and landing flight mode, theaircraft comprising: an airframe including first and second wings withfirst and second pylons extending therebetween; a plurality ofpropulsion assemblies including two propulsion assemblies coupled to thefirst wing and two propulsion assemblies coupled to the second wing,each of the propulsion assemblies including a nacelle and a tailassembly having at least one active aerosurface; and a flight controlsystem operable to independently control each of the propulsionassemblies; wherein, for each of the propulsion assemblies, the tailassembly is rotatable relative to the nacelle such that the activeaerosurface has a first orientation generally parallel to the wings anda second orientation generally perpendicular to the wings.
 2. Theaircraft as recited in claim 1 wherein, each of the propulsionassemblies further comprises an actuator and wherein, for each of thepropulsion assemblies, the actuator is operable to rotate the tailassembly relative to the nacelle.
 3. The aircraft as recited in claim 1wherein, each of the propulsion assemblies further comprises an actuatorand wherein, for each of the propulsion assemblies, the actuator isoperable to translate the tail assembly relative to the nacelle betweena retracted configuration and an extended configuration.
 4. The aircraftas recited in claim 1 wherein, in the first orientation, each of theactive aerosurface further comprises a horizontal stabilizer.
 5. Theaircraft as recited in claim 1 wherein, in the first orientation, eachof the active aerosurfaces further comprises an elevator.
 6. Theaircraft as recited in claim 1 wherein, in the first orientation, eachof the active aerosurfaces provides pitch control to the aircraft. 7.The aircraft as recited in claim 1 wherein, in the second orientation,each of the active aerosurfaces further comprises a vertical stabilizer.8. The aircraft as recited in claim 1 wherein, in the secondorientation, each of the active aerosurfaces further comprises a rudder.9. The aircraft as recited in claim 1 wherein, in the secondorientation, each of the active aerosurfaces provides yaw control to theaircraft.
 10. The aircraft as recited in claim 1 wherein, duringvertical takeoff and landing flight maneuvers, each of the activeaerosurfaces is in the first orientation.
 11. The aircraft as recited inclaim 1 wherein, during hover flight maneuvers, each of the activeaerosurfaces is in the first orientation.
 12. The aircraft as recited inclaim 1 wherein, during transitions from vertical takeoff and landingflight maneuvers to forward flight maneuvers, each of the activeaerosurfaces is in the first orientation.
 13. The aircraft as recited inclaim 1 wherein, during transitions from forward flight maneuvers tovertical takeoff and landing flight maneuvers, each of the activeaerosurfaces is in the first orientation.
 14. The aircraft as recited inclaim 1 wherein, during forward flight maneuvers, each of the activeaerosurfaces is in the second orientation.
 15. An aircraft operable totransition between a forward flight mode and a vertical takeoff andlanding flight mode, the aircraft comprising: an airframe includingfirst and second wings with first and second pylons extendingtherebetween; a plurality of propulsion assemblies including twopropulsion assemblies coupled to the first wing and two propulsionassemblies coupled to the second wing, each of the propulsion assembliesincluding a nacelle, a tail assembly having at least one activeaerosurface, a first actuator and a second actuator; and a flightcontrol system operable to independently control each of the propulsionassemblies, the flight control system including an active aerosurfacecontrol system operable to control operations of each of the first andsecond actuators; wherein, for each of the propulsion assemblies, thefirst actuator is operable to rotate the tail assembly relative to thenacelle such that the active aerosurface has a first orientationgenerally parallel to the wings and a second orientation generallyperpendicular to the wings; and wherein, for each of the propulsionassemblies, the second actuator is operable to translate the tailassembly relative to the nacelle between a retracted configuration andan extended configuration.
 16. The aircraft as recited in claim 15wherein, in the first orientation, each of the active aerosurfacesfurther comprises a horizontal stabilizer operable to provide pitchcontrol to the aircraft.
 17. The aircraft as recited in claim 15wherein, in the first orientation, each of the active aerosurfacesfurther comprises an elevator operable to provide pitch control to theaircraft.
 18. The aircraft as recited in claim 15 wherein, in the secondorientation, each of the active aerosurfaces further comprises avertical stabilizer operable to provide yaw control to the aircraft. 19.The aircraft as recited in claim 15 wherein, in the second orientation,each of the active aerosurfaces further comprises a rudder operable toprovide yaw control to the aircraft.
 20. The aircraft as recited inclaim 15 wherein, during vertical takeoff and landing flight maneuvers,during hover flight maneuvers and during transitions between verticaltakeoff and landing flight maneuvers and forward flight maneuvers, eachof the active aerosurfaces is in the first orientation and wherein,during forward flight maneuvers, each of the active aerosurfaces is inthe second orientation.